Turbine cooling construction

ABSTRACT

A turbine cooling construction for use in a gas turbine engine which permits the use of stoichiometric temperatures within the turbine. More specifically, an inlet turbine temperature in excess of 3,000* F. may be utilized.

lnventors Loren H. White East Hartford, Conn.; David K. Dorer,Rochester, Mich. Appl. No. 780,944 Filed Nov. 29, 1968 Patented Aug. 24,I971 Assignee United Aircraft Corporation East Hartford, Conn.

TURBINE COOLING CONSTRUCTION References Cited UNITED STATES PATENTS3,377,803 4/1968 Prachar .l 60/3966 FOREIGN PATENTS 1 711,985 7/1954Great Britain 60/3966 Primary Examiner-Samuel Feinberg Anomey-Jack N.McCarthy peratures within the turbine. More specifically, an inletturbine temperature in excess of 3,000 F. may be utilized.

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us. ca. 60/39., 60/267,4l5/l77,4l6/95 in. ca. F02c7/12 FieldofSeareh...416/95; 415/177, 178, l75;60/39.66

M A I I i a W i I 1 2 36 I II me i Z 47 i'i H This invention relates togas turbine engines, and more specifically to a cooling construction forthe turbine components of a gas turbine engine.

Fundamentally, gas turbine performance is directly related to thetemperature of the gas entering the turbine. The benefits to be expectedfrom cooling of gas turbine components, both in terms of higher outputand higher efficiency, are so well known as to require no furtherelaboration Many proposals have been directed to this end, particularlyto the cooling of the turbine blades, which are the most critical point,in the turbine wheel. Notwithstanding these schemes, most gas turbinesare kept within endurable temperature limits simply by diluting themotive gases with excess air, which'results in low efficiency. 1

In most of the prior art constructions, turbine disks and buckets havebeen cooled by circulation of air over or through the wheel and throughthe buckets. Since very considerable amounts of air are required for asmall degree of cooling and the air must be forced through against thepressure of the turbine motive fluid, this too is inefficient.

The present invention avoids the penalties to efficiency by employing acooling construction which employs a cryogenic fuel as the coolingfluid. The use of this type system represents a significant advancementin turbine cooling technology since its effectiveness is such thatturbine inlet temperatures, corresponding to stoichiometric fuel-airratios, may be achieved. More specifically, this means that turbineinlet temperatures maybe almost doubled in relation to current enginesand that substantial gains in gas turbine performance can be realized.

SUMMARY OF THE INVENTION The primary object of the present invention isto provide a turbine cooling construction, the effectiveness of which issuch that turbine inlet temperatures corresponding to stoichiometricfuel-air ratios may be achieved, hence providing substantial gains ingas turbine performance.

The present invention accomplishes the foregoing objective by using acooling fluid which is a cryogenic fuel. The coolant is supplied from anexternal source to a first flow path means which is in communicationwith passageways within both the turbine disk and turbine blades;therefore, the coolant flows from the external source through theturbine wheel and blades, hence maintaining the foregoing components ata desired temperature. The actual metal temperatures of the turbine diskand blade are controlled by the coolant velocity, which is a function ofthe flow coefficient through the passageways of the first flow pathmeans, the turbine disk, and the turbine blades.

To provide an overall turbine with a substantially increased efficiency,a third flow path means is provided to the turbine nozzle upstream ofthe turbine disk and blades. The third flow path is in communicationwith a second flow path hereinafter described, and transmits the coolantfluid from the turbine disk area to the turbine nozzle vanes, whichcontain passageways for the flow of the coolant therethrough. It shouldbe clear, therefore, that the coolant cools not only the turbine diskand blades but also the turbine nozzle vanes, the metal temperature ofeach being a function of the flow coefficient of each flow path means.The second flow path means hereinbefore mentioned transmits the fluidfrom the turbine disk exit area to the combustion chamber which isnormally located upstream of the turbine nozzle vane. The combustionchamber contains injector means for injecting the coolant, which is acryogenic fuel, into the combustion chamber and hence provides anadditional source of motive gases to the turbine.

The present invention additionally handles a problem which is generallyassociated with a cooling construction which utilizes a fuel as thecooling medium, the problem generally being one of an increased firehazard inthe' engine. In the preferred embodiment of the presentinventiom'theturbinedisk is ,util. ized'to provide an air compartmentforward or upstream of the disk, while providing a ,fuel compartment aftor downstream of the disk. This siibstantiallyminimizes the tire":hazard within the present construction. m

BRIEF DESCRIPTION O'FTHE pitiAwtNo DESCRIPTION OF THE PREFERREDEMBODIMENT Referring now to the single FIGURE, the turbine coolingconstruction of the invention is shown as applied to a turbojet engineof conventional design, the Savin US. Pat. No. 2,747,367 being a typicalexample. The engine includes a compressor, not shown herein, whichdischarges air into combustion chamber 2. Fuel is suitably supplied tocombustion chamber 2 and the products of combustion are utilized todrive turbine 4.

Positioned around turbine 4 is turbine casing 6 within which ispositioned a supply means for supplying a cooling fluid, hereinafterdescribed in greater detail. The turbine 4 includes a turbine rotorassembly including turbine wheel or disk 8, on which is mounted aplurality of turbine blades 10. The center of the turbine disk 8 isclosed whereby the disk 8 forms a solid wallfacing the forwardcompartment 60. Positioned within turbine disk 8 is a plurality ofpassageways 12, the passageways 12 having appropriate inlets l4 andoutlets 16. The inlets l4 and outlets 16 of the disk open to therearward side of the solid wall formed by the disk. Similarly positionedwithin each of the turbine blades 10 is a plurality of passageways 18,the passageways l2 and I8 being in communication with one another.Extending from supply means 80 in turbine casing 6 is first flow pathmeans 20. As herein illustrated, first flow path means 20 comprises adouble-wall construction, one wall being indicated by the referencecharacter 22 and the second wall by the reference character 24. In thepresent embodiment, first flow path means are shown as passing through apassageway 26 in exit guide vane 28 positioned downstream of the turbine4, it therefore being clear that at least exit guide vanes 28 may becooled by any coolant flowing through the exit guide vanes.

As to the coolant, it has been found desirable from the standpoint ofengine-thrust weight performance to employ a coolant with both a'coolantand a fuel. Genetically, it has been found that a cryogenic fuelsatisfies this criteria. In the cooling construction of the presentinvention the coolant is supplied to the supply means 80 from where itflows through the first flow path means 20, through exit guide vanes 28and hence connectively cools the exit guide vanes 28. The coolantcontinues to flow through first flow path means 20 to turbine disk inlet14 and then internally of turbine disk 8 through passageways l2 andturbine blades 10 through passageways 18. As hereinbefore noted, thepassageways l2 and 18 are in communication and it should be noted thatfor the sake of simplicity only one passageway in the dish, one turbineblade, one passageway in the turbine blade, one exit guide vane and onepassageway in the exit guide vane have been illustrated, and thatobviously in the usual engine construction, this is not the conventionalconstruction.

It has been successfully demonstrated that by causing a cryogenic fuelto flow through first flow path means 20, and the turbine disk 8 andturbine blades 10, as hereinbefore described, that it is possible to runa stoichiometric turbine, or more specifically, a turbine with an inlettemperature in excess of 3000 F. The construction of the first flow pathmeans 20 and other flow path means hereinafter described is ofsignificance in that at the flow coefficient through these flow pathmeans that actually controls or determines the level of the metaltemperature of the turbine disk 8 and turbine blades 10. Morespecifically, metal temperatures throughout the entire turbineconstruction are controlled by the coolant velocity which is controlledas a function of the flow coefficient of the flow paths.

The coolant flow after convectively removing heat from the turbine 4exits from disk 8 as at outlet 16. It enters second flow path means 30which like first flow path 20 comprises a double-wall construction asindicated at 32 and 34. As illustrated in the present embodiment, secondflow path means extend from outlet 16, through turbine casing 6 andpasses forwardly to combustion chamber 2. Combustion chamber 2 includesinjector means 36 through which air from the engine compressor (notshown) and the coolant are injected into combustion chamber 2. It shouldbe clear that the present system by recovering the turbine coolant andutilizing it as a primary fuel significantly reduces engine fuelconsumption and improves engine efficiency.

The cooling construction of the present invention includes one otherflow path means, i.e., a third flow path means 40 which extends from thesecond flow path means 30 through turbine inlet nozzle 42 and intoinjector means 36. As hereinbefore explained, only one nozzle vane 42and coolant passageway 46 are illustrated for purposes of simplicity.The construction of the third flow path means is similar to the firstand second flow path means and comprises a double-wall passageway asindicated by reference characters 48 and 50.

A problem generally associated with a cooling system using fuel as thecoolant is the fire hazard involved. The present embodiment of theinvention minimized the fire hazard by compartmenting the engine withturbine disk 8. More specifically, all engine air remains forward ofdisk 8 as in compartment 60 and any free coolant or cryogenic fuel iscontained aft or downstream of disk 8 as in compartment 62.

We claim:

1. In a gas turbine engine including a combustion section, a turbinesection and an exit section, a double walled outer casing extendingaround said combustion section, turbine section and exit section, saidturbine section including a turbine rotor mounted for rotation therein,said rotor including a rotor disk, a plurality of blades mounted on'saidturbine disk, the center of said rotor disk being closed, said rotordisk thereby forming a solid wall, fixed means extending rearwardly fromsaid disk with its outer wall forming an annular exit passageway withthe casing, inlet means at the rear end of thete xi t section of thedouble walled outer casing for admitting "a cryogenic fluid, exit guidevanes extending between the inner wall of said doublewall casing and theouter wall of the fixed means within said annular exit passageway, saidfixed means being double walled with a central passage means connectingthe rear double wall to a location adjacent the rear of the disk and aconcentric passage means connecting the front double wall to a secondlocation adjacent the rear of the disk, said exit guide vanes andcooperating double walls being divided so that the double walled casingaround said exit section is connected to the rear part of the doublewalled fixed means and the forward part of the double walled fixed meansis connected to the double walled casing around the turbine section andcombustion section, said disk having passageways therein which haveinlets at one location and exits at another, said passageways beingconnected to passages in the blades, means connecting the inlets of saidpassageways to said central passage means and the exits to saidconcentric passage means, a double walled inner casing located radiallyinwardly from the. double walled casing and forwardly of the turbinerotor forming an annular combustion chamber with said double walledouter casing, turbine inlet nozzles extending between the double walledouter casing and the rear end of the double walled inner casing justupstream of the blades, said nozzles having a passageway thercthroughconnecting .the interior of the two double walls, a plurality ofinjection means connecting the double walled outer casing to the forwardend of the double walled inner casing, said injection means havingopenings for injecting a fluid into the combination chamber.

1. In a gas turbine engine including a combustion section, a turbinesection and an exit section, a double walled outer casing extendingaround said combustion section, turbine section and exit section, saidturbine section including a turbine rotor mounted for rotation therein,said rotor including a rotor disk, a plurality of blades mounted on saidturbine disk, the center of said rotor disk being closed, said rotordisk thereby forming a solid wall, fixed means extending rearwardly fromsaid disk with its outer wall forming an annular exit passageway withthe casing, inlet means at the rear end of the exit section of thedouble walled outer casing for admitting a cryogenic fluid, exit guidevanes extending between the inner wall of said double-wall casing andthe outer wall of the fixed means within said annular exit passageway,said fixed means being double walled with a central passage meansconnecting the rear double wall to a location adjacent the rear of thedisk and a concentric passage means connecting the front double wall toa second location adjacent the rear of the disk, said exit guide vanesand cooperating double walls being divided so that the double walledcasing around said exit section is connected to the rear part of thedouble walled fixed means and the forward part of the double walledfixed means is connected to the double walled casing aroUnd the turbinesection and combustion section, said disk having passageways thereinwhich have inlets at one location and exits at another, said passagewaysbeing connected to passages in the blades, means connecting the inletsof said passageways to said central passage means and the exits to saidconcentric passage means, a double walled inner casing located radiallyinwardly from the double walled casing and forwardly of the turbinerotor forming an annular combustion chamber with said double walledouter casing, turbine inlet nozzles extending between the double walledouter casing and the rear end of the double walled inner casing justupstream of the blades, said nozzles having a passageway therethroughconnecting the interior of the two double walls, a plurality ofinjection means connecting the double walled outer casing to the forwardend of the double walled inner casing, said injection means havingopenings for injecting a fluid into the combination chamber.